The use of composite materials in modem vehicles has expanded in recent years inasmuch as composite materials provide a relatively strong structure without the weight penalty that is associated with their metallic counterparts. Composite sandwich panels, which consist of upper and lower skin structures affixed to a core structure, are used extensively as an efficient method for transferring axial, bending, and shear loads. The upper and lower skin structures of the sandwich panel are typically fabricated from fiber reinforced resin matrix material and are designed to carry the axial and bending loads. The core structure is designed to carry the transverse shear loads and is usually a honeycomb structure, which comprises an array of cells defined by an associated plurality of cell walls.
A simple composite sandwich panel is fabricated by compacting and curing the upper and lower skin structures to a single honeycomb core structure in one step. The step of compaction removes air pockets that may be trapped between the skin structures and the core. Compaction is achieved by placing the structures on a mold surface, disposing a vacuum bag over the structures and sealing the bag to the mold, and then applying a vacuum to the assembly. The curing process involves placing the compacted assembly in an autoclave and applying pressure to the assembly while exposing the same to an elevated temperature. In order for the panel to cure correctly, the applied curing pressure must transfer completely through the assembly so as to urge the skin structures to bond to the honeycomb core structure. For a simple sandwich panel, the pressure that is applied to the upper and lower skin structures is reacted by the cell walls of the honeycomb core structure. The process of curing skin structures and honeycomb core structures in one step is known as co-curing and provides a simple yet rigid sandwich panel.
Some structural designs, however, require the incorporation of an internal or intermediate composite laminate of fiber reinforced resin matrix material between two honeycomb core structures. For instance, the RAH-66 Comanche aircraft utilizes lightly loaded sandwich panels on many external surfaces. These panels are designed to carry or react the airloads which are applied to the aircraft during flight and, therefore, a single honeycomb panel typically provides sufficient strength. However, certain portions of the panel require additional stiffness in order to support locally high loads which are applied, e.g., walking loads, attachment loads, etc. Referring to FIG. 1a, additional stiffness may be provided by bonding a second honeycomb structure directly to the first structure. This is achieved either by separately curing the panels then bonding them together or, more preferably, by co-curing the panels together in a single step.
Additional stiffness may also be provided by locally increasing the thickness of the skin structures. FIG. 1b illustrates an internal attachment point on a honeycomb panel. The primary deficiency with this arrangement is that the applied fitting load, F.sub.fitting, transfers to the core as a peel load. The allowable peel strength of a honeycomb panel is relatively low as compared to its shear strength and, therefore, such an attachment would not be adequate. FIG. 1c illustrates an improved internal attachment point. The applied fitting load is reacted by the inner and outer skin structures and transfers to the core as shear. The main deficiency with this type of design is that the external surface will no longer be smooth, thus adversely affecting the low observable characteristics of the airframe as well as the aerodynamic flow along the airframe surface. In order to alleviate these concerns, an additional honeycomb panel is affixed on the external surface over the attachment point. The additional honeycomb panel may be cured separately from the first honeycomb structure then bonded thereto, however, it is more preferable to co-cure the two panels in a single step.
The foregoing complex composite sandwich structures have, to date, been difficult to manufacture using a co-curing process inasmuch as the precise location of the first and second honeycomb structures is required to adequately transfer the applied curing pressure through the cell walls to the intermediate composite laminate. FIG. 1d illustrates the problem associated with improper placement of the honeycomb structures. The intermediate laminate is not shown in order to illustrate that the curing pressure will only transfer to the intermediate laminate as point loads at the intersection 10 of the cell walls of the upper and lower honeycomb core structures 12,14. This type of loading does not adequately pressurize the intermediate laminate during curing, hence, yielding a deficient part. Precise placement of honeycomb structures on the intermediate composite laminate is an exceedingly difficult process requiring exact trimming of the honeycomb core. Moreover, proper placement of the honeycomb structures prior to curing does not guarantee that the honeycomb core structures will not shift during the actual curing process. Additionally, if the honeycomb core is trimmed incorrectly or has been damaged during the fabrication process such that a depression is formed thereon, the applied curing pressure, P.sub.applied, will transfer to only one side of the intermediate laminate, resulting in a `bridging` of the depression as shown in FIGS. 1e,f.
A need therefore exists for an improved co-curing process for fabricating a composite sandwich structure having an intermediate composite laminate disposed between two honeycomb core structures.